EP2674677A2 - Ensemble de refroidissement pour une chemise de chambre de combustion pour système de turbine à gaz - Google Patents

Ensemble de refroidissement pour une chemise de chambre de combustion pour système de turbine à gaz Download PDF

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Publication number
EP2674677A2
EP2674677A2 EP13171620.1A EP13171620A EP2674677A2 EP 2674677 A2 EP2674677 A2 EP 2674677A2 EP 13171620 A EP13171620 A EP 13171620A EP 2674677 A2 EP2674677 A2 EP 2674677A2
Authority
EP
European Patent Office
Prior art keywords
flow
sleeve
combustor liner
impingement
cooling assembly
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Withdrawn
Application number
EP13171620.1A
Other languages
German (de)
English (en)
Inventor
Venugopal Polisetty
Sridhar Venkat Kodukulla
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
General Electric Co
Original Assignee
General Electric Co
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by General Electric Co filed Critical General Electric Co
Publication of EP2674677A2 publication Critical patent/EP2674677A2/fr
Withdrawn legal-status Critical Current

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/002—Wall structures
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/005—Combined with pressure or heat exchangers
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F23R3/04—Air inlet arrangements
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
    • F23R2900/03043—Convection cooled combustion chamber walls with means for guiding the cooling air flow
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
    • F23R2900/03044—Impingement cooled combustion chamber walls or subassemblies

Definitions

  • the subject matter disclosed herein relates to gas turbine systems, and more particularly to a combustor liner cooling assembly.
  • a combustor section of a gas turbine system typically includes a combustor chamber disposed relatively adjacent a transition piece, where a hot gas passes from the combustor chamber through the transition piece to a turbine section. At least a portion of the combustor chamber is often surrounded by a flow sleeve, while at least a portion of the transition piece is surrounded by an impingement sleeve.
  • the flow sleeve typically includes a plurality of apertures for providing impingement cooing for portions of a liner of the combustor. An additional airflow passes from a region defined by the impingement sleeve and the transition piece to a region defined by the flow sleeve and the combustor liner.
  • the impingement cooling of the liner of the combustor is achieved by cooling jets that are pushed onto the liner in a direction relatively perpendicular to the additional airflow flowing from the region proximate the impingement sleeve to the region proximate the flow sleeve.
  • the additional airflow often disrupts the cooling jets, thereby resulting in reduced cooling efficiency.
  • a combustor liner cooling assembly for a gas turbine system includes a combustor liner defining a combustor chamber. Also included is a flow sleeve surrounding at least a portion of the combustor liner, wherein the flow sleeve includes at least one aperture row comprising a plurality of apertures, each of the plurality of apertures impinging a cooling flow jet onto the combustor liner.
  • a plurality of flow redirecting components disposed proximate an aft end of the flow sleeve, wherein the plurality of flow redirecting components divert an impingement cross-flow flowing relatively perpendicular to the cooling flow jet, thereby providing the cooling flow jet an undisturbed flow path to the combustor liner.
  • a combustor liner cooling assembly for a gas turbine system includes a combustor liner defining a combustor chamber. Also included is a flow sleeve surrounding at least a portion of the combustor liner and having an aft end, wherein the flow sleeve includes a plurality of apertures for impinging a plurality of cooling flow jets onto the combustor liner.
  • an impingement sleeve disposed proximate the aft end of the flow sleeve, wherein an impingement flow path is defined by the impingement sleeve and a transition duct, wherein an impingement cross-flow flows through the impingement flow path into a region between the flow sleeve and the combustor liner.
  • a plurality of flow redirecting components disposed proximate the aft end of the flow sleeve, wherein the plurality of flow redirecting components divert the impingement cross-flow.
  • a combustor liner cooling assembly for a gas turbine system includes a combustor liner defining a combustor chamber. Also included is a flow sleeve surrounding at least a portion of the combustor liner and having an aft end, wherein the flow sleeve includes a plurality of aperture rows, wherein each of the plurality of aperture rows comprises a plurality of apertures extending circumferentially around the flow sleeve, wherein each of the plurality of apertures impinges a cooling flow jet onto the combustor liner.
  • each of the plurality of flow redirecting components is circumferentially aligned with a corresponding first row aperture for diverting an impingement cross-flow entering a region between the flow sleeve and the combustor liner proximate the aft end of the flow sleeve.
  • FIG. 1 partial schematic illustrates a combustor section of a gas turbine system and is referred to generally with numeral 10.
  • the combustor section 10 includes a transition piece 12 having a transition duct 14 at least partially surrounded by an impingement sleeve 16 disposed radially outwardly of the transition duct 14. Upstream thereof, proximate a forward end 18 of the impingement sleeve 16 is a combustor liner 20 defining a combustor chamber 22.
  • the combustor liner 20 is at least partially surrounded by a flow sleeve 24 disposed radially outwardly of the combustor liner 20.
  • a forward sleeve 26 is located at the junction between the forward end 18 of the impingement sleeve 16 and an aft end 28 of the flow sleeve 24.
  • the combustor section 10 uses a combustible liquid and/or gas fuel, such as a natural gas or a hydrogen rich synthetic gas, to run the gas turbine system.
  • the combustor chamber 22 is configured to receive and/or provide an air-fuel mixture, thereby causing a combustion that creates a hot pressurized exhaust gas.
  • the combustor chamber 22 directs the hot pressurized gas through the transition piece 12 into the turbine section (not illustrated), causing rotation of the turbine section.
  • the presence of the hot pressurized exhaust gas increases the temperature of the combustor liner 20 surrounding the combustor chamber 22, particularly proximate a downstream end 30 of the combustor liner 20.
  • a plurality of apertures 32 within the flow sleeve 24 are arranged to provide impinged air in the form of a plurality of cooling jets 34 onto the combustor liner 20.
  • the plurality of apertures 32 may optionally include "thimbles" (not illustrated) which protrude toward the combustor liner 20, providing an enclosed region to deliver the plurality of cooling jets 34 toward the combustor liner 20.
  • An impingement cross-flow 36 flows relatively perpendicularly to the plurality of cooling jets 34 and provides a convective cooling effect on the combustor liner 20 while flowing from downstream to upstream along the combustor liner 20. Specifically, the impingement cross-flow 36 flows from a region defined by the impingement sleeve 16 and the transition duct 14 to a region defined by the flow sleeve 24 and the combustor liner 20.
  • FIG. 2 an enlarged view of the aft end 28 of the flow sleeve 24, the forward sleeve 26 and the forward end 18 of the impingement sleeve 16 is shown in greater detail.
  • the plurality of apertures 32 within the flow sleeve 24 may be arranged in one or more circumferential rows proximate the aft end 28 of the flow sleeve 24.
  • the forward sleeve 26 includes at least one, but typically a plurality of flow redirecting components 38 operably coupled thereto that are disposed along an inner surface of the forward sleeve 26 in a circumferentially spaced arrangement.
  • the plurality of flow redirecting components 38 may be integrally formed with the forward sleeve 24 or may be fastened thereto.
  • Each of the plurality of flow redirecting components 38 includes a flow redirecting surface 40 that is arranged to interact with the impingement cross-flow 36 that is flowing upstream toward the combustor liner 20 and the flow sleeve 24.
  • Each of the plurality of flow redirecting components 38 is relatively circumferentially aligned with at least one of the plurality of apertures 32.
  • the plurality of flow redirecting components 38 are described above and illustrated as being operably coupled to the forward sleeve 26, it is contemplated that alternative embodiments may include operable coupling of the plurality of flow redirecting components 38 to the impingement sleeve 16 proximate the forward end 18 thereof. Additionally, it is contemplated that the plurality of flow redirecting components 38 may be operably coupled to the aft end 28 of the flow sleeve 24, provided that the plurality of flow redirecting components 38 are disposed downstream of the plurality of apertures 32.
  • a first embodiment of the plurality of flow redirecting components 38 comprises a semi-circular geometry, with the flow redirecting surface 40 arranged to interact with the impingement cross-flow 36, as described above. As the impingement cross-flow 36 interacts with the flow redirecting surface 40, the impingement cross-flow 36 is diverted around the flow redirecting surface 40. As noted above, the plurality of flow redirecting components 38 are relatively aligned with the plurality of apertures 32, and therefore also the plurality of cooling jets 34 flowing relatively perpendicularly to the impingement cross-flow 36.
  • a second embodiment of the plurality of flow redirecting components 38 is shown and is similar in construction to that of the first embodiment illustrated in FIGS. 3-5 .
  • the second embodiment of the plurality of flow redirecting components 38 includes a plurality of holes 42 for reducing the formation of vortices upon recirculation of the impingement cross-flow 36 subsequent to passing the flow redirecting surface 40.
  • a third embodiment of the plurality of flow redirecting components 38 is illustrated and is similar in construction to the embodiments described above.
  • the third embodiment of the plurality of flow redirecting components 38 includes a first portion 44 having the previously described semi-circular geometry, which includes the flow redirecting surface 40 terminating in a first end 46 and a second end 48. Extending axially upstream from at least one of the first end 46 and the second end 48 is a second portion 50 that provides additional axial structure for the impingement cross-flow 36 to flow along.
  • the additional structure provided by the second portion 50 reduces the axial space between the plurality of flow redirecting components 38 and the plurality of cooling jets 34, thereby reducing the likelihood of the impingement cross-flow 36 disrupting the plurality of cooling jets 34.
  • the third embodiment is illustrated with the plurality of holes 42 described above in relation to the second embodiment, however, it is to be appreciated that the third embodiment may include the second portion 50, but not the plurality of holes 42.
  • FIGS. 8-10 additional embodiments of the plurality of flow redirecting components 38 are illustrated.
  • the additional embodiments are similar to the embodiments described above, but rather than a semi-circular geometry, the additional embodiments include a triangular geometry.
  • a fourth embodiment ( FIG. 8 ) of the plurality of flow redirecting components 38 includes a triangular geometry having a flow redirecting peak 52 arranged to interact with the impingement cross-flow 36, as described above with respect to the flow redirecting surface 40 of the semi-circular embodiments.
  • a fifth embodiment ( FIG. 9 ) includes the plurality of holes 42.
  • a sixth embodiment ( FIG. 10 ) includes a first triangular portion 54 extending from the flow redirecting peak 52 to a first end 56 and a second end 58, where at least one second portion 60 may extend therefrom, similar to the third embodiment described above.
  • the sixth embodiment may include the at least one second portion 60, but not the plurality of holes 42.
  • the plurality of flow redirecting components 38 are described above as having particular geometric shapes, however, it is to be understood that any suitable geometric shape capable of diverting the impingement cross-flow 36 may be employed as the plurality of flow redirecting components 38.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP13171620.1A 2012-06-13 2013-06-12 Ensemble de refroidissement pour une chemise de chambre de combustion pour système de turbine à gaz Withdrawn EP2674677A2 (fr)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US13/495,674 US20130333388A1 (en) 2012-06-13 2012-06-13 Combustor liner cooling assembly for a gas turbine system

Publications (1)

Publication Number Publication Date
EP2674677A2 true EP2674677A2 (fr) 2013-12-18

Family

ID=48578905

Family Applications (1)

Application Number Title Priority Date Filing Date
EP13171620.1A Withdrawn EP2674677A2 (fr) 2012-06-13 2013-06-12 Ensemble de refroidissement pour une chemise de chambre de combustion pour système de turbine à gaz

Country Status (5)

Country Link
US (1) US20130333388A1 (fr)
EP (1) EP2674677A2 (fr)
JP (1) JP2013256950A (fr)
CN (1) CN103486615A (fr)
RU (1) RU2013126601A (fr)

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2865850B1 (fr) * 2013-10-24 2018-01-03 Ansaldo Energia Switzerland AG Agencement de refroidissement par impact

Families Citing this family (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JP6202976B2 (ja) 2013-10-10 2017-09-27 三菱日立パワーシステムズ株式会社 ガスタービン燃焼器
KR101557453B1 (ko) 2014-01-15 2015-10-06 두산중공업 주식회사 가스터빈의 이중벽 슬리브 냉각구조를 구비한 라이너 및 그 냉각방법
EP2955442A1 (fr) * 2014-06-11 2015-12-16 Alstom Technology Ltd Agencement de paroi refroidie par convection
WO2016013585A1 (fr) * 2014-07-25 2016-01-28 三菱日立パワーシステムズ株式会社 Cylindre de chambre de combustion, chambre de combustion, et turbine à gaz
EP3205937B1 (fr) * 2016-02-09 2021-03-31 Ansaldo Energia IP UK Limited Agencement de paroi refroidie par impact
DE102016224632A1 (de) * 2016-12-09 2018-06-14 Rolls-Royce Deutschland Ltd & Co Kg Plattenförmiges Bauteil einer Gasturbine sowie Verfahren zu dessen Herstellung
KR102051988B1 (ko) * 2018-03-28 2019-12-04 두산중공업 주식회사 이중관 라이너 내부 유동가이드를 포함하는 가스 터빈 엔진의 연소기, 및 이를 포함하는 가스터빈
WO2021167001A1 (fr) * 2020-02-20 2021-08-26 川崎重工業株式会社 Structure de refroidissement de bride pour un moteur à turbine à gaz
US11470748B1 (en) 2020-03-09 2022-10-11 Smart Wires Inc. Liquid cooling of high current devices in power flow control systems

Family Cites Families (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20090255268A1 (en) * 2008-04-11 2009-10-15 General Electric Company Divergent cooling thimbles for combustor liners and related method

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
None

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2865850B1 (fr) * 2013-10-24 2018-01-03 Ansaldo Energia Switzerland AG Agencement de refroidissement par impact
US9970355B2 (en) 2013-10-24 2018-05-15 Ansaldo Energia Switzerland AG Impingement cooling arrangement

Also Published As

Publication number Publication date
CN103486615A (zh) 2014-01-01
JP2013256950A (ja) 2013-12-26
US20130333388A1 (en) 2013-12-19
RU2013126601A (ru) 2014-12-20

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